This invention relates to a new category of turbojet engines and the application of turbojet engines to aircraft and other vehicles. The turbojet engines are of a type in which the bypass fan blades are integrated with the compressor in association with the turbine blades that are surrounded by an annular combustion chamber. The turbojet engines of this invention are improvements over the engines described in my U.S. Pat. No. 4,845,941 entitled “Gas Turbine Engine Operating Process”, issued Jul. 11, 1989, my U.S. Pat. No. 5,003,766 entitled “Gas Turbine Engine”, issued Apr. 2, 1991, my U.S. Pat. No. 5,177,954 entitled “Gas Turbine Engine With Cooled Turbine Blades”, issued Jan. 12, 1993, and in my U.S. Pat. No. 5,341,636 entitled “Gas Turbine Engine Operating Method”, issued Aug. 30, 1994.
Conventional turbojet engines are comprised of separated modules, including bypass ducted fans, axial and centrifugal compressors, combustion chambers, and gas turbines. When assembled along a common axis the modules and components of the conventional turbojet engines combine to form an elongated engine that lacks the compactness required for many of the applications described herein.
The bypass fan is the main propulsion module and all these other modules and components cooperate to finally drive this module. The complexity of modern turbo jets has reached the maximum level, and the cost is beyond any limit of affordability by the majority of the world.
The thermal efficiency of conventional turbojets is limited to 30% at full loads and drops to near 10% at part loads.
The power density is limited by the maximum temperature of the combustion. At 25% of the maximum stoichiometric level, the air fuel ratio is 60/1, instead of a stoichiometric 15/1, making all turbojet engines at least four times larger than an engine operating at the stoichiometric level, with a resulting lower efficiency and greater expense.
For military applications, specifically drone airplanes and cruise missiles, the high cost for aircraft designed to be lost in combat is a major expense for defense and a burden on the national economy.
It is a primary object of this invention to provide a turbojet engine of high efficiency that combines isothermic compression of a part of the intake air in hollow fan blades for stoichiometric combustion with the bypass air flowing between the fan blades providing cooling for the compressed air directed to the combustion chamber.
This continuation-in-part application includes additional configurations of turbojet engines of the type described in my patent application Ser. No. 10/292,829 filed on Nov. 12, 2002. The subject embodiments of the turbojet engine in this disclosure are designed for high altitude commercial and military aircraft in both atmospheric and space flight. The common feature of the two embodiments of the high altitude turbojet engines is the use of rocket propulsion with atmospheric oxygen in atmospheric flight and enriched liquid oxygen at high altitude and space flight. These hybrid systems provide a universal propulsion system for a variety of military and commercial applications with a tremendous reduction in the costs of operation.
The greatest barriers for high performance gas turbines and jet engines are the limitations on maximum temperature acceptable for combustion gases and the limitations in the pressure ratio of air compression in the engine cycle. The metallurgical properties of gas turbine blades and the limited cooling schemes available for gas turbine blades combine to severely limit turbine inlet temperatures. In turn this requires a high air-fuel ratio of 50/1 to 60/1 to maintain inlet temperatures within the range acceptable for modern turbine blade designs. The embodiments of the turbojet engines having rocket capabilities solves these problems and produce the maximum absolute thermodynamic performance for aircraft in atmospheric and space flight.